The invention relates to an apparatus and method for assessing the dynamic stiffness capability of a hydraulic actuator, and more particularly, to a method of assessing the capability of a hydraulic flight control actuator to provide sufficient dynamic stiffness for limiting control surface flutter while installed on an aircraft.
In typical flight-by-wire flight control systems, fluid driven flight control actuators are widely known and used. These actuators can be linear or rotary devices, and typically include at least one cylinder divided into two variable volume chambers by a piston. In conventional systems, the piston supports a piston rod that is attached to an aircraft control surface. Fluid pressure can be selectively applied to one or both chambers through a servo valve or other similar valve to drive the piston for repositioning the control surface to which the piston rod is attached.
An important property of flight control actuators is their dynamic stiffness capabilities, i.e., the ability of the fluid in the actuator to react to external loads to resist movement of the control surface. That is, the dynamic stiffness of a flight control actuator is the ability of the actuator to react to the inertial load generated by unwanted attempts to accelerate the control surface. And, more importantly, control surface flutter is prevented by the ability of the flight control actuator to provide a stiff connection between the control surface and the aircraft structure. Thus, it is highly desirable in aircraft systems to monitor the stiffness capability of flight control actuators, as failures within the actuator can result in a significant reduction in the dynamic stiffness of the actuator.
Rather than assessing an actuator""s dynamic stiffness capabilities directly, aircraft in service today perform automated tests to measure a related parameter and then infer from this data an actuator""s ability to provide sufficient stiffness. For instance, in flight by wire control systems the dynamic stiffness of an actuator can be inferred based on the damping force capability of the actuator. For a given actuator design, the manufacturer determines an acceptable range of damping force values over which the actuator can normally operate. If the actuator is determined to operate within the acceptable damping force range, the dynamic stiffness of the actuator is presumed to be acceptable.
The damping force (Fdamp) provided by an actuator equals the differential pressure (DP) across the actuator piston multiplied by the piston area (Ap), Fdamp=DP*Ap. For a given actuator design, the piston area (Ap), the area of the damping orifice (Aorifice), and the flow constant (Kq) determined by fluid and orifice properties are often known, and the flow rate (Q) of the hydraulic fluid through the damping orifice is equal to the rate at which the piston is moving multiplied by the piston area (Ap), Q=(Aorifice)*Kq(DP)1/2. Solving for DP:
DP=Q2/(Aorifce*Kq)2 
Fdamp=Ap*DP
Fdamp=Aorifice*Kq 
Fdamp=Ap*(Rate*Ap)2/(Aorifice*Kq)2.
Thus, the damping force is proportional to the rate squared.
The above referenced equation and analysis is performed on an actuator while in the damping mode. It has always been assumed that if the actuator performed satisfactorily in the damping mode according to the above formula, the actuator was also acceptable in the active mode. However, analysis data shows that a small leakage path across the actuator in the active mode, while resulting in small changes in the damping force (Fdamp), may result in significant changes in the dynamic stiffness. Thus, it may be desirable to assess directly the dynamic stiffness capability of a particular actuator to limit the flutter of a particular aircraft control surface.
The invention relates to an apparatus and method for verifying the stiffness capabilities of a flight control actuator. The apparatus comprises known sensors, electronics and hydraulic components.
A method for assessing the dynamic stiffness of a flight control actuator using the apparatus of the present invention can include the steps of (1) providing a control surface actuator system including an actuator having a portion thereof coupled to an aircraft control surface; (2) applying a command signal to the actuator control system, causing the production of a pressure differential across the actuator in response to the command signal based on the ability of the actuator to react to an inertial load generated by the attempted acceleration of the control surface; and (3) processing a pressure signal representative of the pressure differential produced across the actuator. The method can further include the step of comparing the pressure signal to known pressure differential values produced under similar conditions for specific actuator and control surface combinations.